Discrete co-flow jet (DCFJ) airfoil

ABSTRACT

The present invention provides an aircraft having one or more fixed wings in a flying wing configuration, where the aircraft further includes a high performance co-flow jet (CFJ) circulating about at least a portion of an aircraft surface to produce both lift and thrust.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation-in-part-of U.S. utility applicationSer. No. 12/119,193, filed May 12, 2008, entitled CO-FLOW JET AIRCRAFT,which is a continuation-in-part-of PCT Application Ser. No.PCT/US07/10122, filed Apr. 24, 2007, entitled EMISSIONLESS SILENT ANDULTRA-EFFICIENT AIRPLANE USING CFJ AIRFOIL, which is related to andclaims priority to U.S. provisional patent application Ser. No.60/796,042, filed Apr. 28, 2006, entitled ENGINELESS EMISSIONLESS SILENTAND ULTRA-EFFICIENT AIRPLANE USING CFJ AIRFOIL; this application is alsoa continuation-in-part-of U.S. utility patent application Ser. No.11/064,053, filed Feb. 23, 2005, entitled HIGH PERFORMANCE AIRFOIL WITHCO-FLOW JET FLOW CONTROL, which is related to and claims priority toU.S. provisional patent application Ser. No. 60/603,212, filed Aug. 20,2004, entitled HIGH PERFORMANCE AIRFOIL WITH CO-FLOW JET FLOW CONTROL,the entirety of all of which is incorporated herein by reference.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

n/a

FIELD OF THE INVENTION

The present invention relates to aircraft design, propulsion, andoperation.

BACKGROUND OF THE INVENTION

Conventional aircraft have traditionally made use of propellers or jetengine propulsion systems to generate thrust and the wings, in turn,generate the lift necessary to support the weight of the aircraft. Thesetwo systems, the propulsion and lift-generating systems, have alwaysbeen treated separately. Unlike man-made vehicles, birds, insects andother flying animals do not have separate propulsion and lift systems.They rely on flapping wings to generate both lift and thrust. The downstroke of the flapping wings has a very large angle of attack (AoA) tothe relative flow. Vortex shedding at both leading and trailing edges isthe dominant flow phenomenon of a bird flapping its wings. The result isthat the dynamic circulation of the flapping wing is so high that itgenerates sufficient lift to support the body weight of a bird, and atthe same time, the high circulation generates very strong low pressuresuction at the wing leading edge that results in a net thrust.Ornithopters use the same principle to fly, however, they are generallylimited to very small unmanned air vehicles (UAV). This is generally dueto the fact that driving the flapping wings for large aircraft is verydifficult and inefficient. From studying bird flight, it can be deducedthat if the circulation is sufficiently high, a wing can generate bothlift and thrust. In view of the above, it would be desirable to providean aircraft having an integrated propulsion and lift generating system,thereby reducing aircraft complexity, and greatly increasing performanceand efficiency.

SUMMARY OF THE INVENTION

The present invention advantageously provides a system for an aircrafthaving an integrated propulsion and lift generating system and a methodof operation, thereby reducing aircraft complexity, and greatlyincreasing performance and efficiency. In particular, the presentinvention may provide an aircraft having one or more fixed wings in aflying wing configuration, where the aircraft further includes a highperformance co-flow jet (CFJ) airfoil to produce both lift and thrustrather than a conventional propulsion system (i.e., a propeller or jetengine). As a result, the energy expenditure is significantly reducedcompared to that of a conventionally powered aircraft, as the energyconsumption is largely limited to the power to provide a fluid flowacross a portion of the aircraft, which does not necessarily require acombustion device. In addition, the maneuverability and safety of theaircraft is further enhanced due to the increased stall margin of theCFJ airfoil.

For this aircraft, the co-flow jet airfoil produces both the lift andthrust. The concept of the CFJ airfoil may generate extraordinaryperformance with a net zero drag (for cruise) or a net negative drag(thrust, for acceleration), as well as extremely high lift and stallmargin. The aircraft may include a flying wing design with an increasedsurface area about which the CFJ may be integrated. By using such aconfiguration, the CFJ airfoil may extend across a substantial portionof the fuselage section of the aircraft.

The aircraft of the present invention may be advantageous for use acrossa wide range of applications. For example, the aircraft and methods ofoperation of the present invention may include an unmannedreconnaissance aircraft, small personal aircraft, commercial airliners,and many other applications.

The aircraft of the present invention may not necessarily be limited toflight on Earth, but also for exploratory missions to other planets. Forexample, the CFJ airplane may be particularly well suited for flight inthe Martian atmosphere due to reduced energy consumption, enhancedmaneuverability and safety, extremely short take off/landing distance,soft landing and take off with very low stall velocity. Such performanceis desirable due to the limited amount of fuel that can be carried in amission to Mars, the limited availability of take-off and landing space,as well as the challenges of flying in a low density atmosphere in alaminar flow regime.

The present invention can be also applied to helicopter or rotorcraftrotor blades to remove rotor blade dynamic stall, increase rotor bladelift, thrust, and efficiency. The CFJ or Discrete CFJ airfoil can beapplied to any fluid-machineries that use airfoils. These includeaircraft, gas turbine engine turbomachinery, helicopter/rotorcraft,pumps, wind turbines, propellers, among other blade related machinery

The present invention further provides an aircraft having an aircraftbody defining a leading edge, a trailing edge, a first wing having afirst wingtip and a second wing having a second wingtip. The aircraftmay include an injection opening proximate the leading edge; a recoveryopening located between the injection opening and the trailing edge; oneor more engine portions positioned between the recovery opening and thetrailing edge., where the one or more engine portions include an exhaustportion, compressor stages, and/or an air inlet leading to thecompressor stages.

The aircraft body may define a shape substantially similar to a flyingwing, wherein the injection opening and/or the recovery opening mayextend along a substantial length between the first and second wingtips.In addition, the recovery opening may be in fluid communication with afluid intake of the one or more engines, and may also be in fluidcommunication with a fluid path leading to the injection opening. Theone or more engines can provide a pressurized fluid output to theinjection opening. The aircraft body may define an arcuate tail portionhaving an upper surface, where one or more engines are at leastpartially integrated into the upper surface of the arcuate tail portion,and the arcuate tail portion may define one or more depressions adjacentan exhaust path of the one or more engines. One or more ailerons mayreside on the tail portion, while the first and second wings may defineunitary bodies devoid of movable flaps.

BRIEF DESCRIPTION OF THE DRAWINGS

A more complete understanding of the present invention, and theattendant advantages and features thereof, will be more readilyunderstood by reference to the following detailed description whenconsidered in conjunction with the accompanying drawings wherein:

FIG. 1 shows an embodiment of a co-flow jet airfoil in accordance withthe present invention;

FIG. 2 depicts a fluid flow field for a conventional airfoil of theprior art;

FIG. 3 illustrates a fluid flow field for an embodiment of a co-flow jetairfoil in accordance with the present invention;

FIG. 4 is a graphical illustration of a comparison of a measured liftcoefficient for both conventional airfoils of the prior art as well asembodiments of a co-flow jet airfoil in accordance with the presentinvention;

FIG. 5 is a graphical illustration of a measured injection momentumcoefficient for embodiments of a co-flow jet airfoil in accordance withthe present invention;

FIG. 6 is a graphical illustration of a comparison of a measured dragpolar for both conventional airfoils of the prior art as well asembodiments of a co-flow jet airfoil in accordance with the presentinvention;

FIG. 7 is an additional graphical illustration of a comparison of ameasured drag polar for both conventional airfoils of the prior art aswell as embodiments of a co-flow jet airfoil in accordance with thepresent invention;

FIG. 8 shows an embodiment of an aircraft in accordance with the presentinvention;

FIG. 9 depicts a baseline NACA 6425 airfoil of the prior art;

FIG. 10 is a graphical illustration of three-dimensional streamlines atroot for an embodiment of an aircraft having an angle of attack equal to40° in accordance with the present invention;

FIG. 11 is a graphical illustration of a three-dimensional coefficientof lift versus angle of attack for an embodiment of an aircraft inaccordance with the present invention;

FIG. 12 is a graphical illustration of a comparison of a measured dragpolar for both a conventional airfoil of the prior art as well asembodiments of a co-flow jet airfoil in accordance with the presentinvention;

FIG. 13 is an illustration of three dimensional surface pressurecontours at an angle of attack equal to 0° for an embodiment of anaircraft in accordance with the present invention;

FIG. 14 is a graphical illustration of the momentum coefficient Cμversus angle of attack for an embodiment of an aircraft in accordancewith the present invention;

FIG. 15 is a graphical illustration of the lift and drag profile at anangle of attack equal to 0° along a wing span of an embodiment of anaircraft in accordance with the present invention;

FIG. 16 is a graphical illustration of a three-dimensional wake profileplot for an embodiment of an aircraft having an angle of attack equal to0° in accordance with the present invention;

FIG. 17 is a perspective view of an embodiment of an aircraftconstructed in accordance with the principles of the present invention;

FIG. 18 is another perspective view of an embodiment of an aircraftconstructed in accordance with the principles of the present invention;

FIG. 19 is a top view of an embodiment of an aircraft constructed inaccordance with the principles of the present invention;

FIG. 20 is a side view of an embodiment of an aircraft constructed inaccordance with the principles of the present invention;

FIG. 21 is a front view of an embodiment of an aircraft constructed inaccordance with the principles of the present invention;

FIG. 22 is an additional front view of an embodiment of an aircraftconstructed in accordance with the principles of the present invention;

FIG. 23 is a rear perspective view of an embodiment of an aircraftconstructed in accordance with the principles of the present invention;

FIG. 24 is a cross-sectional schematic of an embodiment of an aircraftconstructed in accordance with the principles of the present invention;

FIG. 25 is a perspective view of an embodiment of an aircraftconstructed in constructed in accordance with the principles of thepresent invention;

FIG. 26 is a side perspective view of the engagement of the tabs withthe aircraft shown in FIG. 25; and

FIG. 27 is a top perspective view of the aircraft shown in FIG. 25.

DETAILED DESCRIPTION OF THE INVENTION

It will be appreciated by persons skilled in the art that the presentinvention is not limited to what has been particularly shown anddescribed herein above. In addition, unless mention was made above tothe contrary, it should be noted that all of the accompanying drawingsare not to scale. A variety of modifications and variations are possiblein light of the above teachings without departing from the scope andspirit of the invention, which is limited only by the following claims.

The present invention advantageously provides an aircraft having anintegrated propulsion and lift generating system, thereby reducingaircraft complexity, and greatly increasing performance and efficiency.In particular, the present invention may provide an aircraft having oneor more fixed wings in a flying wing configuration, where the aircraftfurther includes a high performance co-flow jet (CFJ) circulating aboutat least a portion of an aircraft surface to produce both lift andthrust rather than or in addition to a conventional propulsion system(i.e., a propeller or jet engine). The present invention provides anaircraft having an injection slot near a leading edge of the aircraftbody and a recovery slot near a trailing edge of the aircraft body. Ahigh energy jet or fluid stream may be injected near the leading edge inthe same direction of the main fluid flow across the aircraft andsubstantially the same amount of mass flow may be recovered near thetrailing edge. The fluid flow jet may then be recirculated to maintain azero-net mass flux flow control.

Now referring to FIG. 1, an aerodynamic structure 10 is shown having achord length, a leading edge 14, and a trailing edge 16. The leadingedge 14 is the portion of the aerodynamic structure 10 which interactswith fluid first, i.e., the “front” of the structure 10, with thetrailing edge 16 located at the rear point of the aerodynamic structure10. The aerodynamic structure 10 further includes a first surface 18that generally defines a surface extending from the leading edge 14 tothe trailing edge 16. A second surface 20, which is opposite the firstairfoil surface 18, also generally defines a surface extending from theleading edge 14 to the trailing edge 16. The first surface 18corresponds to the suction side of the aerodynamic structure 10, i.e.,the first surface 18 experiences a pressure lower than that experiencedacross the second surface 20 when the aerodynamic structure 10 issubjected to a fluid flow.

The first surface 18 also defines an injection opening 22 locatedbetween the leading edge 14 and the trailing edge 16, and furtherdefines a recovery opening 24 located in between the injection opening22 and the trailing edge 16. In an exemplary embodiment, the injectionopening 22 is located less than 25% of the chord length from the leadingedge 14 of the structure. To increase the effectiveness of the adversepressure gradient to enhance mixing, the injection opening 24 may belocated downstream of the leading edge suction peak of the aircraft bodyand/or airfoil. However, the benefits of the present invention may berealized with the injection opening located within 80% of the chordlength from the leading edge 14. Moreover, the recovery opening 24 maybe located less than 25% of the chord length from the trailing edge 16of the aerodynamic structure. Similarly to the injection openingplacement, however, the benefits of the present invention may berealized with the recovery opening 24 located within 80% of the chordlength from the trailing edge 16. The injection opening 22 defines aninjection opening height, which may have a value that is generally lessthan 5% of the chord length. The recovery opening 24 defines a recoveryopening height, which may have a value generally less than 5% of thechord length. While the injection and recovery openings illustrated havea fixed size, an alternative embodiment can include openings capable ofhaving their height varied through the use of mechanical means in whichat least a portion of the first surface 18 is moveable, thereby changingthe height of either the injection opening or the recovery opening.

The aircraft may include a pressurized fluid source and/or a vacuumsource 26. The vacuum source may provide a pressure lower than anambient pressure. The pressurized fluid source may be in fluidcommunication with the injection opening 22, and can include a pump orother means of pressurizing a fluid. The vacuum source may be in fluidcommunication with the recovery opening 24, and may also include apumping apparatus.

An exemplary use of the CFJ aircraft provides a method for reducing theboundary layer separation of an aerodynamic structure. The aerodynamicstructure 10 may be operated such that a first mass of fluid is routedfrom the pressurized fluid source towards the injection opening. Thefirst mass may be routed by any means of conducting a fluid, i.e., aconduit, tubing, or the like. The first mass may be dispersed out of theinjection opening and directed substantially tangent to the exteriorsurface of the aerodynamic structure 10 and towards the recovery opening24. Concurrently, the vacuum source creates a pressure at the recoveryopening 24 lower than that of the environment external to the recoveryopening 24, resulting in a second mass of fluid being drawn into therecovery opening 24.

Although the injection and recovery of fluid along the aerodynamicstructure can be realized by separate and independent injection andrecovery resources, the fluid flow can also be recirculated by a pumpsystem. The second mass can then be drawn into the recovery opening 24and directed to the front stage of a compressor or the inlet where thepressure is low. The fluid flow is hence recirculated to save energyexpenditure. Further, the fluid can be provided by a compressed airsupply, such as a pressurized tank.

The fundamental mechanism involved with a CFJ aircraft structure is thatthe severe adverse pressure gradient on the suction surface of anaircraft and/or airfoil strongly augments the turbulent shear layermixing between the main fluid flow and the jet. The mixing then createsthe lateral transport of energy from the jet to the main flow and allowsthe main flow to overcome the large adverse pressure gradient and remainattached even at high angles of attack. The stall margin is hencesignificantly increased. At the same time, the high momentum jetdrastically increases the circulation about the aircraft, whichsignificantly augments lift, reduces drag and/or even generates thrust(net negative drag). For example, as shown in FIGS. 2 and 3, a typicalcomparison illustrates that the prior art airfoil has significantseparation at high angle of attack, whereas the CFJ airfoil has fluidflow that is significantly more attached at the trailing edge. The flowbegins to separate at high angles of attack for a conventional baselineairfoil with no flow control. With the implementation of the CFJ,however, the Co-Flow Jet over the suction surface of the airfoil acts toovercome the adverse pressure gradient, thus keeping the flow attached,even at extremely large angles of attack. With an attached flow at ahigh angle of attack, the stall margin and maximum lift aresignificantly increased. For example, experimental wind tunnel testshave shown that flow will remain attached for certain airfoils at anangle of attack as high as 43 degrees, which comprises a significantincrease over existing methods and structures. The energized boundarylayer from the CFJ also results in drag reduction and even thrustgeneration (net negative drag) at low angles of attack. Continuing torefer to FIGS. 2 and 3, the flow over the lifting surface is shown asseparated at approximately the midline of the surface of the baselineairfoil, while at this same angle of attack, the streamlines depictedaround the Co-Flow Jet airfoil show the flow remains attached.

When compared with a conventional, prior art circulation control (CC)airfoil, the working mechanism of aircraft of the present invention isdifferent. A CC airfoil relies on large leading edge (LE) or trailingedge (TE) to have the Coanda effect and enhance circulation. The largeTE or LE may generate large drag during cruise. The aircraft of thepresent invention may include wall jet mixing to energize the main flowand overcome the adverse pressure gradient so that the flow can inducehigh circulation and remain attached at high AoA. The CC airfoil dumpsaway the jet mass flow, which is a considerable penalty to thepropulsion system. The aircraft of the present invention, on the otherhand, may recirculate the jet mass flow and achieve the zero net jetmass flux to have very low energy expenditure. Compared with thesynthetic jet flow control, the enhancement of aircraft and/or airfoilperformance having an injected and recovered fluid flow stream or jet ismuch more drastic because the interaction of the main flow with thesynthetic jet generated either by acoustic waves or plasma is generallytoo weak. As a result, the aircraft of the present invention maysimultaneously achieve three radical improvements at low energyexpenditure: lift enhancement, stall margin increase, and drag reductionor thrust generation.

Control volume analysis indicates that the drag or thrust of a CFJairfoil measured in the wind tunnel is the actual force acting on theairfoil or aircraft system in the stream-wise direction. This is not thesame as the CC airfoil, which must consider the equivalent drag due tothe suction penalty from the free-stream. For a CC airfoil, theequivalent drag is significantly larger than the drag measured in a windtunnel and is also substantially larger than the drag of a CFJ airfoil.For a CFJ airfoil, the suction penalty is already included in themeasured drag and is off set by the higher circulation and strongerleading edge suction induced by the CFJ. The drag reduction mechanism ofa CFJ airfoil is not based on the conventional concept to reduce theskin friction. Instead, it relies on the inclusion of the pressureresultant force, which overwhelms the skin friction. When the leadingedge suction is very strong, the low pressure at the leading edgeprovides a resultant force that is forward-pointing and is greater thanthe skin friction, resulting in the production of thrust. When a thrustis generated by the wing, the need for a conventional engine orpropulsion system is significantly reduced, if not eliminatedaltogether.

Experimental analysis has indicated the performance enhancementsprovided by the CFJ airfoil. For example, an airfoil having a 6″ chordand 12″ span (to fit the 12″×12″×24″ test section) was analyzed. Due tothe small size of the tested airfoil, a thick airfoil, NACA0025 , wasselected to facilitate the instrumentation and internal ducts. Thefreestream Mach number was 0.1, the Reynolds number was about 4×10⁵. Tomimic the turbulent boundary layer with a large Reynolds number as in arealistic flight situation, a leading edge trip was implemented toenforce the turbulent boundary layer. The co-flow jet airfoils are namedusing the following convention: CFJ4dig-INJ-REC, where 4dig is the sameas NACA 4 digit convention, INJ is replaced by the percentage of theinjection slot size to the chord length and REC is replaced by thepercentage of the recovery slot size to the chord length. For example,the CFJ0025-065-196 airfoil has the injection slot height of 0.65% ofthe chord and the recovery slot height of 1.96% of the chord. TheCFJ0025-131-196 has a twice as large injection slot size with the samerecovery slot size as the CFJ0025-065-196. The injection and recoveryslot of the tested airfoil were located at 7.11% and 83.18% of the chordfrom the leading edge.

FIG. 4 illustrates the comparison of measured lift coefficient for thebaseline NACA0025 airfoil and the CFJ0025-065-196 airfoil with theinjection total pressure coefficient given (the last number in thelegend, normalized by freestream total pressure). During a test, theinjection total pressure was held as constant while the AoA varies. Ahigher injection total pressure will yield a higher injection momentumcoefficient, and hence a higher lift coefficient and stall margin. Thebottom two curves with circle and cross symbols are for the baselineNACA0025 airfoil with and without LE trip. It shows that the one withtrip delays stall by about 4 degrees of AoA. This is because the fullyturbulent boundary layer with the trip is more resistant to flowseparation. The very bottom curve is the CFJ airfoil without the jet on.It has less stall AoA than the baseline airfoil because the injectionand suction slot steps weaken the boundary layer and make separationoccur at a smaller AoA. The results shown in FIG. 3 indicate that theCFJ airfoil significantly increases the lift and stall angle of attackover that of a conventional airfoil.

Table 1 lists the aerodynamic parameters of the baseline NACA0025airfoil and the CFJ0025-065-196 airfoil with injection total pressurecoefficient of 1.27. Table 1 indicates that the C_(Lmax) of theCFJ0025-065-196 airfoil is 5.04, whereas the maximum lift coefficient ofthe baseline airfoil is 1.57, a 220% increase. The baseline airfoilstalls at AoA of 19°, while the CFJ0025-065-196 airfoil stalls atAoA=44°, an increase of 153%.

TABLE 1 C_(Dmin) AoA_(CL) = Cμ_(CL) = (AoA = Airfoil 0 0 AoA_(CLmax)C_(Lmax) Cμ_(CLmax) 0°) Base- 0° 0.0 19° 1.57 0.0 0.128 line NACA- 0025CFJ0- -4° 0.187 44° 5.04 0.28 -0.036 025- 065-196

FIG. 5 illustrates the injection momentum coefficient, C_(μ), of theCFJ0025-065-196 airfoil at three different injection total pressures.The injection mass flow rate and velocity are determined by theinjection total pressure and the mainflow static pressure at theinjection location. The injection total pressure is held constant whilethe AoA varies. When the AoA is increased, the LE suction is strongerand hence the local static pressure at the injection location decreases.The injection velocity therefore will increase, so will the mass flowrate and the momentum coefficient as shown in FIG. 4. For the highestinjection total pressure coefficient of 1.27, the momentum coefficientvaries from 0.184 to 0.3. The lowest injection total pressurecoefficient of 1.04 has the momentum coefficient varying from 0.05 to0.1, which increases the C_(Lmax) by 113% and AoA stall margin by 100%.These results indicate that even the small momentum coefficient is veryeffective to enhance the lift and stall margin.

FIG. 6 is the drag polar of the CFJ0025-065-196 airfoil. The dragcoefficient of the CFJ airfoil is significantly reduced and has a smallregion of negative drag (thrust). For example, at C_(L)=1, forC_(μ)=0:071, the CFJ0025-065-196 airfoil drag reduction is 19%; forC_(μ)=0:197, the drag reduction is 90%, which increases the L/D tenfold.At lower C_(L) values with the total pressure coefficient of 1.27, thedrag reduction is over 100% because the drag is negative and becomesthrust. When the drag becomes zero or negative, the conventionalaerodynamic efficiency measurement of L/D may not be meaningful since itapproaches infinity. At low AoA, the CFJ airfoil wake is filled with theenergized mainflow and has inversed velocity deficit. In this case, theairfoil has no drag, but thrust. The airfoil drag can be decomposed totwo parts: skin friction and pressure drag. The skin friction drag doesnot vary much when the AoA changes. Rather, it is the large pressureresultant force that significantly decreases the total drag or generatesthrust (negative drag). The strong leading edge suction makes asignificant contribution to the thrust generation or drag reduction.

FIG. 7 is the drag polar of the CFJ0025-131-196 airfoil with theinjection size and jet mass flow rate about twice larger than those ofthe CFJ0025-065-196 airfoil. The thrust (negative drag) region of theCFJ0025-131-196 airfoil is significantly larger than that of theCFJ0025-065-196 airfoil. For example, the CFJ0025-131-196 airfoil stillhas thrust at C_(L)=2:0 , whereas the CFJ0025-065-196 airfoil has nothrust when the C_(L) is greater than 0.9. This means that when the jetmass flow rate is increased, the thrust is also increased. In general,we can achieve the required thrust and lift by adjusting the jetstrength. Based on the mass conservation law, a suction is neededwhenever an injection is used for an airfoil flow control. For a CFJairfoil, the suction occurs on the airfoil suction surface near trailingedge. For a CC airfoil, the suction occurs by drawing the jet mass flowfrom freestream through the engine inlet.

A control volume analysis was conducted to analyze the lift and dragbreakdowns due to the jet ducts and the CFJ airfoil performance with andwithout jet suction. Based on the momentum and mass equations, the dragof a CFJ airfoil is:D═R′ _(x) −F _(xcfj)═∫_(wake) ρV _(e)(V _(∞) −V _(e))δγ  Eq. (1)

where, R′_(x) is the CFJ airfoil surface pressure and shear stressintegral in x-direction. F_(xcfj) is the reaction force generated by theinjection and suction jet ducts in x-direction. V_(e) is the velocitydownstream of the airfoil. Eq.(1) indicates that the drag measured inthe wind tunnel is the actual drag that the 2D CFJ airfoil will be actedon. The suction penalty is already included in F_(xcfj) and is off setby the higher circulation and stronger leading edge suction induced bythe CFJ that is included in R′_(x). The integral in Eq.(1) shows thatthe drag of a CFJ airfoil is equal to the drag calculated by the wakeprofile. This is the same as a conventional non-controlled airfoil.However, this is not true for a circulation control airfoil. When CCairfoil is used for an airplane, the actual drag, or the “equivalentdrag”, needs to add the penalty caused by drawing the jet mass flow fromfreestream. The equivalent drag coefficient of a CC airfoil can bewritten as:C _(Dequiv) ═C _(Dwindtunnel) +Cμ(V _(ei) /V _(j))+Cμ(V _(ei) /V _(j) γM² _(ei))   Eq. (2)

The first term on the right hand side of Eq.(2) is the CC airfoil dragmeasured in wind tunnel, the 2nd term is the ram drag, and the 3rd termis the drag due to the captured area. The subscript ei stands for engineinlet, j stands for injection jet. The results based on CFD simulationindicate that the equivalent drag of a CC airfoil is also significantlylarger than the drag measured in a wind tunnel and is substantiallygreater than the drag of a CFJ airfoil. The power consumed by a CCairfoil is hence also significantly more.

It has been suggested that the suction occurring on the airfoil suctionsurface such as the CFJ airfoil is much more beneficial to enhanceairfoil performance than having the suction from freestream such as theCC airfoil. Compared with the airfoil with injection only, the CFJairfoil has higher lift, higher stall margin, lower drag, and lowerpower required. A concept study based on CFD simulation indicates thatit is possible for the CFJ airfoil to exceed the inviscid limit ofmaximum lift coefficient due to the high jet velocity inducing highcirculation of the airfoil. For a cambered CFJ airfoil modified fromNACA0025, a CFD calculated lift coefficient of 9.7 is obtained withoutusing any flap, which is far greater than the inviscid maximum liftcoefficient limit of 7.8.

In summary, the CFJ airfoil concept provides the followingadvantages: 1) significantly enhanced lift and suppress separation; 2)drastically reduced drag or generated thrust; 3) significantly increasedAoA operating ranges and stall margins; 4) substantially reduced energyexpenditure; 5) equally applicable to airfoils of varying thickness; 6)controllable for an entire flight and/or any portion thereof; 7) can beused for low and high speed aircraft; 8) easy implementation with nomoving parts; 8) equally applicable for both fixed wings and rotatingrotor blades.

The concept of CFJ airfoil has demonstrated the extraordinaryperformance to enhance lift, generate thrust, and increase stall margin.Now turning to FIG. 8, an aircraft 28 of the present invention may havean injection slot 30 and a recovery slot 32 such that a co-flow jet orstream of fluid is circulated across a substantial portion of a surfaceof the aircraft. The aircraft 28 defines an upper surface, where theupper surface may include a first portion and a second portion. Thesecond portion may be recessed with respect to the first portion, andthe second portion may comprise a substantial amount of the surface areaof the upper surface. For example, the recessed portion may extend froma location proximate or otherwise close to a first wing tip to alocation proximate a wing tip on the opposite side of the aircraft.

By providing the concept of the CFJ across an increased surface area ofthe aircraft 28, the aircraft thus would have a reduced and/oreliminated need for the inclusion of a propeller or jet engine systembecause the CFJ airfoil itself is capable of generating thrust.Accordingly, the thrust generated by implementation of the CFJ conceptacross a substantial portion of the aircraft surface can overcome the3-D induced drag due to tip vortices. In particular, the aircraft 28 ofthe present invention may include a flying wing configuration as fluidjet can flow across almost the entire aircraft surface to achieve themaximum benefit, resulting in the generation of lift and thrust whereverit is applied. Thus, the only drag that needs to be overcome by the CFJairfoil thrust would be the induced drag due to tip vortices. In orderto operate, the airplane 28 may include a pumping system to draw the jetmass flow near the trailing edge and inject the jet near the leadingedge as illustrated in FIG. 1. In addition, at different phases of theflight mission, the lift and thrust can be controlled by adjusting thejet strength. For example, during take off, a stronger jet may be usedto generate high thrust and high lift, while at cruising speed, a weakerjet may be used due to lower lift coefficient and the amount of thrustrequired to remain in flight. Upon landing, the jet velocity and/or massmay be adjusted to allow the aircraft to fly at high angle of attackwith high lift and high drag.

A conventional airplane draws the air flow from the free-streamenvironment through the engine inlet, energizes the air through thecombustion process, and then exhausts the high momentum air to theenvironment through the engine nozzle. Such a process is purely forthrust generation and has no interaction with the wing. The energytransfer from the chemical energy of combustion to mechanical energy(momentum increase) is usually very inefficient and accompanies a verylarge thermal energy (total enthalpy) loss of 50% or more. A CFJ wingdraws the air flow on the suction surface of the wing near the trailingedge, pressurizes the air within the wing and then exhausts the same airnear the wing leading edge. Such a process has a direct interaction withthe wing and enhances the wing lift by inducing a large circulation andgenerates a thrust at the same time. The mass flow of the jet may besubstantially less than that of a jet engine. The jet recirculation orpumping process (recovery and injection) requires less power than thatof a jet engine and can be achieved using electric power. The energytransfer is from mechanical energy (pumping the CFJ) to mechanicalenergy (high momentum injection jet) and therefore the efficiency ismuch higher. No combustion process is needed and as such, emissions maybe completely eliminated.

The power required to pump the jet for this aircraft may besignificantly less than the power required to run a conventional jetengine. When the power is consumed to generate the CFJ and enhance lift,it also reduces the drag or produces thrust at the same time. For theconventional airplanes, the power system is used only to overcome thedrag without enhancing lift coefficient. The equivalent L/D of the CFJairplane hence is much higher than that of the conventional airplane.Since the lift coefficient of the CFJ airfoil element is significantlyhigher than the conventional airfoil, the overall lifting surface areato have the same payload will thus be much smaller. The weight of theairplane and the drag due to the whetted surface will be alsosignificantly reduced. With no aircraft engines, the weight of theengines and the drag due to the engine nacelles and captured area willalso be removed. The reduced weight and drag will further reduce theenergy consumption.

The power required to pump the jet is determined by the ratio of thetotal pressure at the injection and suction and the mass flow rate ofthe jet. Compared with a jet engine system, the reduction of powerneeded for a CFJ system results from the following: 1) the mass flowrate of the jet may be much smaller than the mass flow rate of the jetengine; the conservative estimation is that the maximum jet mass flowrate would not exceed 30% of that of a conventional jet engine; 2) thetotal pressure ratio to pump the jet may be much smaller than the totalpressure ratio of a jet engine compressor. For example, if the injectiontotal pressure is 2 times the static pressure in the injection slotarea, the injection jet Mach number will be 1.05. Usually, the injectionjet speed will be limited to lower than sonic speed for subsonic flight.Both FIGS. 6 and 7 list the injection total pressure normalized by thefreestream pressure (the last numbers in the legend). They are notgreater than 1.27. The compressor total pressure ratio of a modern jetengine is usually about 30, which is far greater than the total pressurerequired to pump the jet; 3) the CFJ injection and recovery are at themost energy efficient locations. The recovery is near the trailing edgewhere the pressure is the highest on the airfoil except the LEstagnation point. The flying wing embodiment of the aircraft of thepresent invention may also eliminate the need for a conventional tailstructure. Instead, winglets could be located at the wingtips forlateral stability and control. These winglets make use of a symmetricairfoil cross-section. The use of a conventional tail may be avoided inorder to reduce instabilities introduced during planetary entry.Horizontal stability and control may be provided by a more conventionalpair of flaperons on the aft of the wings injection is right downstreamof the leading edge suction peak where the pressure is the lowest. Thepressure gradient is favorable to recirculate the jet and minimize thepower required to pump and energize the jet.; 4) No combustion is neededand hence no thermal loss occurs; 5) The overall engineless airplaneweight and drag is much less than the conventional airplane. The energyexpenditure is hence greatly reduced.

As discussed, use of the CFJ system across at least a portion of anaircraft may significantly reduce energy expenditure. The reduction ofthe power required for an Engineless CFJ airplane could be up to 70% ormore when compared to that of a conventional, combustion driven jetengine. The lower power consumption of a CFJ airplane provides muchlonger range and endurance than a conventional airplane. In addition toenergy expenditure, the CFJ aircraft may have extremely short takeoff/landing (ESTOL) distance due to the very high maximum liftcoefficient. For the same reason, the stall velocity will besignificantly lower than the conventional airplane. The lower stallvelocity will allow soft landing and take off at substantially lowerspeed. Another important use of CFJ airfoil during take off/landing isto enhance the subsonic performance of a supersonic wing for asupersonic airplane. Moreover, due to the high stall margin, the CFJairplane will have significantly higher maneuverability and safetymargin to resist severe weather conditions, such as unexpected gusts ofwind. The high stall margin is also particular useful for Mars airplanesto resist flow separation and stall at low Reynolds number.

Again referring to FIG. 8, unlike most conventional aircraft, where thewings and fuselage are separate structures, the aircraft of the presentinvention may include an airframe where both of these components areincorporated into a single, blended body. This is typically called aflying wing configuration, because the entire aircraft effectively actsas a wing. Because the fuselage may have the same airfoil cross sectionas the wings, it acts as an extension of the same and thus producesadditional lift. This particular embodiment also allows for an increasedcoverage area for the CFJ mechanism, therefore increasing the benefitsgained from using it.

A flying wing design also allows for a reduction in the wingspan of theaircraft. As the fuselage surface is no longer “wasted”, but made toproduce lift, the aircraft can produce increased lift with a shorterwingspan. This feature is desirable particularly for Martianapplications because, in order to reach Mars, the aircraft must bepackaged within an aeroshell. The goal is generally to be able to fitthe aircraft within an aeroshell while minimizing the number of foldsnecessary. An aircraft which needs to unfold once it is deployed intothe atmosphere is generally less stable and safe due to the increasedcomplexity. An increased number of folds will also increase theprobability of failure during deployment, which is the most criticalstep during the aircraft's mission.

As discussed above, the baseline airfoil chosen for comparison to theCFJ aircraft of the present invention is the NACA 6425 airfoil, whichcan be seen in FIG. 9. This airfoil has a camber of 6% located 40% fromthe leading edge, with a maximum thickness of 25% of the chord. Thisairfoil was chosen for its moderate camber and high thickness. The highthickness would allow for comfortable placement of all of the CFJcomponents, such as the pump and ducting. Also, airfoils with highthicknesses will produce higher lift as long as the air flow remainsattached. Conventional aircraft shy away from thicknesses higher than15% due to separation. However, with the use of the CFJ, a higherthickness airfoil can be used without fear of separation occurring, andtherefore an even higher lift can be achieved. A moderate camber waschosen in order to reduce the effect of wing-tip vortices. A highercamber airfoil will produce a higher lift, but there is a penalty in theform of stronger induced drag from wing-tip vortices.

A Computational Fluid Dynamics (CFD) study has been performed for theCFJ aircraft, which demonstrates its increased performance over aconventional aircraft. CFD analysis was performed for both thetwo-dimensional and three-dimensional cases at a range of angles ofattack (AoA) using both the baseline and CFJ airfoil. The simulationswere run at a Reynolds number of 2×10⁶ and a Mach number of 0:1. Thesecomputations were carried out for an aircraft with an aspect ration ofAR=4.

For the 2-D case, the simulations show that separation occurs for thebaseline airfoil at 16° AoA, while flow separation (stall) occurs at 35°AoA for the CFJ airfoil, a 19° difference, as shown in Table 2. Thisconstitutes a significant increase in performance because a higher liftcan be produced without the danger of stalling, even at a relatively lowMach number of 0.1. These results imply that the stall velocity for suchan aircraft would be drastically reduced, and operational angle ofattack vastly increased. A lower stall velocity and increased lift canlead to reduced take-off and landing distances, which is a very highlydesirable trait.

TABLE 2 AoA BL C_(l) CFJ C_(l) BL C_(d) CFJ C_(d) 0 0.0542 2.8517 0.0232−0.9855 10 1.3946 3.9734 0.0408 −0.5939 15 1.5225 5.0729 0.0558 −0.316820 1.4431 5.4402 0.0686 −0.1217 30 1.1147 6.5638 0.1690 0.2613 35 0.93485.5526 0.2342 0.1913

Furthermore, it can be seen that the two-dimensional drag coefficientC_(D) is negative in the case of the CFJ airfoil at angles of attack ashigh as 20°. The drag coefficient becomes positive at high angles ofattack because the form drag has become large enough at that point toovercome the thrust generated by the CFJ airfoil. However, it would beimprobable that the aircraft would ever need to fly in conditions wherethe angle of attack were so high. Even at high angles of attack,however, the drag coefficient of the CFJ airfoil is much lower than thatof the baseline airfoil, reducing the high drag generated at suchconditions.

3-D CFD simulations have been performed for the three-dimensionalEngineless CFJ aircraft in a range of angles of attack from 0° to 45°,using the same Reynolds and Mach numbers as in the 2-D case. The resultsobtained from the post-processing of the data were corrected to includethe jet effects. Although the corrected results are not as favorable asthe uncorrected ones, they still show a very significant increase inperformance over the baseline.

These computations show that flow separates at a very high angle ofattack, about 35° as seen in FIGS. 10 and 11. As can be seen in FIG. 12,the 3-D drag coefficient remains negative within a range of angles ofattack of about −5° to 5°. After this point, the form drag is largeenough to offset the thrust produced by the CFJ airfoil. As can be seenfrom FIG. 13, the pressure drag is greatest in the leading edge of theaircraft (red areas), near the flow stagnation point. This form dragincreases with angle of attack as the profile area presented to theincoming flow becomes larger. This result is slightly lower than the 2-Dcase, but even when the C_(D) becomes positive, it is stillsignificantly lower than that of the baseline case. The C_(D) willprobably be even lower and remain negative at a higher angle of attackfor configurations with a higher aspect ratio, due to the decrease ininduced drag from wing-tip vortices, which is a significant source ofdrag. The momentum coefficient Cμ remains relatively constant throughouta range of angles of attack, as can be seem from FIG. 14.

Wake profile plots for the aircraft at different sections along the wingshow that the drag is more highly negative at the root of the aircraft,and becomes positive towards the wingtips, where induced drag becomessignificant. This can be seen from FIG. 15. However, when averaged overthe wingspan, the net drag is negative at low angles of attack.Normally, the wake of a wing features flow that is slower than insurrounding areas. However, the CFJ wake is particular in that the flowthere is dramatically faster than in surrounding areas, as can see fromFIG. 16. As mentioned before, this type of wake profile will lead to anet thrust being produced.

Now referring to FIGS. 17-22, an aircraft 100 constructed in accordancewith the principles of the present invention is illustrated. Inparticular, aircraft 100 may have an injection slot 102 proximate aleading edge 103 of the aircraft 100, and a recovery slot 104 near thetrailing edge 105 of the aircraft 100 such that a co-flow jet or streamof fluid is circulated across a substantial portion of a surface of theaircraft. The aircraft 100 defines an upper surface 106, where the uppersurface 106 may include a first portion and a second portion. The secondportion may be recessed with respect to the first portion, and thesecond portion may comprise a substantial amount of the surface area ofthe upper surface. For example, the recessed portion may extend from alocation proximate or otherwise close to a first wing tip 108 of theaircraft 100 to a location proximate a wing tip 108′ on the oppositeside of the aircraft. As discussed above, by providing a CFJ across anincreased surface area of the aircraft 100, the aircraft thus hasreduced output requirements for a propeller or jet engine system becausethe CFJ airfoil/fuselage of the aircraft 100 itself is capable ofgenerating thrust.

Continuing to refer to FIGS. 17-22, the aircraft 100 of the presentinvention may include a configuration to enable a fluid jet to flowacross almost the entire upper aircraft surface to achieve the maximumbenefit, resulting in the generation of lift and thrust wherever it isapplied. In particular, the aircraft 100 may generally include a pair ofwings 110, 110′ positioned towards the rear or trailing edge 105 of theaircraft. An elongated, curved forward nose portion 112 may extend fromthe general aircraft body, and a pronounced, curved tail portion 114 mayalso constitute a component of the aircraft 100. The aircraft 100generally includes an airframe where the fuselage and wings areincorporated into a single, blended body. This is typically called aflying wing configuration, because the entire aircraft effectively actsas a wing. Because the fuselage may have a similar airfoil cross sectionas the wings, it acts as an extension of the same and thus producesadditional lift. This particular embodiment also allows for an increasedcoverage area for the CFJ mechanism, therefore increasing the benefitsgained from using it. A flying wing design further allows for areduction in the wingspan of the aircraft. As the fuselage surfacegenerates lift, the aircraft can produce increased lift with a shorterwingspan.

Now referring to FIGS. 23-24, the airplane 100 may include a pumpingsystem and/or thrust components to draw the jet mass flow near thetrailing edge and inject the jet near the leading edge. For example, oneor more engines, turbofans and/or thrust-generation components 116 maybe included on a rear portion of the aircraft 100. The engines mayinclude air intake portions, compressor stages or portions, combustionstages, and an exhaust which may be situated or otherwise positionedwithin a portion of the aircraft body. The engines 116 may be located inbetween the recovery slot 104 and the trailing edge 105 of the aircraft100. As such, the engines 116 may receive air or fluid intake directlyfrom the recovery slot 104 for subsequent combustion for thrustgeneration. In addition, while part of the fluid flowing into therecovery slot 104 is directed towards the engines, a portion may berecirculated or otherwise directed to a reservoir and/or pump 118 forinjection through the injection slot 102 to provide the CFJ operation. Afluid path 119 may provide fluid communication and/or a route forpressurized fluid to flow from the recovery opening 104 and/or the oneor more engines 118 to the injection opening 102.

Referring to FIG. 23 in particular, the engines may be ‘buried’ orotherwise recessed or integrated into a portion of the aircraft body toreduce noise radiation, provide improved aerodynamic characteristics,reduce heat signature, and mask turbofan blades from radar. Inparticular, the engines may be mounted within a top surface of theaircraft on the curved tail portion 114, where the curved tail portion114 defines a contoured, cylindrical surface or depression 120complementary and adjacent to the exhaust portion of each of the one ormore engines 116. Such a configuration can funnel or otherwise directthe exhaust gases towards the trailing edge of the plane withoutdisturbing the fluid flow across the upper surface of the aircraft, andfurther causes most of the engine jet noise to propagate up away fromthe ground, which greatly reduces the noise footprint of the aircraft100 during takeoff and landing.

A conventional aircraft design has protruding engine nacelles thatcontribute to the total drag. In the present aircraft 100, with theengines buried deep in the aft or tail section, drag is greatly reduced.Also, although the high-temperature exhaust from the engines is close tothe trailing edge of the aircraft, the structure of the aircraft issubstantially protected from potentially damaging temperatures fromcooling bypass air exiting the engines to form an encompassing cylinderof cooling air around the hotter exhaust gases. This also greatlyreduces the IR signature of the aircraft.

Furthermore, the suction of the intake and the exhaust jets of theburied engines in the rear part of the airplane enhance the overallcirculation of the flying wing configuration of aircraft 100 andenergize the boundary layer. The recovery slot 104, which is also theengine intake, may run across approximately 80% or more of the entirewingspan. In addition, the slots may be perpendicular to the localairfoil surface to enhance fluid flow and recovery. In a particularexample, the injection slot 102 may include a height of betweenapproximately 0.15% to 0.55% of the mean aerodynamic chord and may beplaced between approximately the 2.5% to 8% chord point on the topsurface of the airfoil. The recovery slot 104 may be positioned betweenapproximately the 65% to 90% chord point, and have a height of betweenapproximately 0.25% and 1% of the mean aerodynamic chord. The angle forthe injection slot may also be between −10 to −20 degrees, while therecovery slot 104 may be between approximately 4 and 10 degrees whenmeasured clockwise from the vertical.

As one or more engines may be located on either side of an aircraftcenterline, asymmetric thrust from the engines could potentially causethe aircraft to yaw. To compensate, the aircraft 100 may further includea split elevator (or elevon pair) 122, 122′ located behind the enginesthat will cause the pitching moments and rolling about the longitudinalaxis of the airplane. The aircraft 100 may further include a controlsystem (not shown) that employs feedback from the propulsion componentsto control the amount of yaw needed to return the aircraft 100 tocoordinated flight at any instant while also allowing a constant liftforce from the CFJ.

Since the CFJ of the aircraft 100 can generate very high lift, the flapand slat systems found in conventional aircraft may be omitted. Withoutthe conventional high lift system, noise during takeoff and landing willbe significantly reduced, and moreover, the number of moving parts maybe significantly reduced, thereby reducing the overall complexity of theaircraft and reducing the likelihood of mechanical failure and the like.Since the mixing effect occurs on the upper surface of the airplane,noise produced by the CFJ will radiate upward. Moreover, a reduced stallspeed causes a further decrease in the noise produced from the wake overthe aircraft at takeoff and landing.

As discussed above, typical combustion engines are typically veryinefficient and have a very large thermal energy (total enthalpy) lossof 50% or more. A CFJ system draws the air flow on the suction surfaceof the wing and/or fuselage near the trailing edge, pressurizes the airwithin the wing and then exhausts the same air near the wing leadingedge. Such a process has a direct interaction with the wing and enhancesthe wing lift by inducing a large circulation and generates a thrust atthe same time.

Use of the CFJ system across at least a portion of an aircraft maysignificantly reduce energy expenditure. In short, the power requiredfor the aircraft 100 is significantly less than that of a conventionalairplane. The ultimate efficiency of an aircraft is determined by theratio of lift to drag. When a part of the power is consumed to generatethe CFJ and enhance lift, it will also reduce the drag, or producethrust at a low angle of attack. For a conventional airplane, the powersystem is used only to overcome the drag and has no interaction with theairframe system to enhance lift. The equivalent L/D of the CFJ airplane100, hence, will be significantly higher than that of the conventionalairplane. Since the lift coefficient of the CFJ airfoil element ishigher than the conventional airfoil, the overall lifting surface areato have the same payload will thus be smaller. The weight of theairplane and the drag due to the wetted surface will be alsosignificantly reduced. With the buried aircraft engines, the drag due tothe engine nacelles will also be removed. As a result, the reducedweight and drag will further reduce the energy consumption.

The power consumed by the CFJ pump alone is:

$P_{pump} = {\frac{\overset{.}{m}\; C_{p}T_{o\; 1}}{\eta}\left( {\left\lbrack \frac{p_{o\; 1}}{p_{o\; 2}} \right\rbrack^{\frac{\gamma - 1}{\gamma}} - 1} \right)}$Where, {dot over (m)}is the CFJ mass flow rate, and; T_(o) and p_(o) arethe total temperature and total pressure, respectively. C_(p) is thespecific heat capacity at constant pressure, γ is the ratio of specificheats (taken to be 1.4), and η is the pump efficiency. Based on theabove equation, the power required to pump or otherwise circulate thejet or fluid stream of the CFJ is dependent on the ratio of the totalpressure at the injection and suction and the mass flow rate of the jet.The CFJ mass flow rate is usually significantly smaller than the enginemass flow rate. The total pressure ratio is also small. For a cruisingexample, the CFJ mass flow rate is 9.7% of the engine mass flow rate. Attakeoff, the CFJ mass flow rate may be approximately 16.9% of the engineflow rate. Hence, the overall power required to pump CFJ, and thusoperated the aircraft 100, is small.

To consider the energy consumption due to the pumping power, anequivalent drag term may be defined that is the summation of the powerrequired to overcome the drag and the CFJ pumping power divided by theflight velocity as the following:

$D_{eq} = \frac{{DV}_{\infty} + P_{pump}}{V_{\infty}}$

The equivalent drag is used to determine QUEIA's L/D for aerodynamicefficiency, which is critical to determine the range and endurance. Inthis exemplary design, the total pressure ratio was taken as 1.1 and theefficiency is taken as 80%. The equivalent lift-to-drag ratio obtainedfor aircraft 100 is thus 39.4. Such a high LID results in low energyexpenditure and extended operating range.

The takeoff/landing distances and the stall velocity are primarilydetermined by the maximum lift coefficient. The stall velocity ofaircraft 100 is significantly lower than a conventional airplane due tothe very high maximum lift coefficient. Consequently, QUEIA will displayshort takeoff and landing (“STOL”) performance. The decreased stallvelocity will reduce runway distance use, and hence increase the airportcapacity. In addition, at different phases of the flight mission, thelift and thrust can be controlled by adjusting the jet strength. Forexample, during take off, a stronger jet may be used to generate highthrust and high lift, while at cruising speed, a weaker jet may be useddue to lower lift coefficient and the amount of thrust required toremain in flight. Upon landing, the jet velocity and/or mass may beadjusted to allow the aircraft to fly at high angle of attack with highlift and high drag.

A significant reduction in noise emission may be achieved through thetakeoff and landing performance alone. The steep climb and descentangles of 8 and −12 degrees that the aerodynamic performance of the CFJpermits ensures that the aircraft climbs more efficiently, or travelshigher for a given horizontal distance. Also, when the aircraft 100descends for landing, the descent may begin closer to the airport toreduce the noise footprint on the airport's surroundings. The glidingperformance rendered by the extremely large L/D value also allowsdescents with minimum power settings that are very efficient and silent

Overall, the present invention provides an advanced aircraft system thattightly integrates the airframe, Co-Flow Jet flow control airfoil, andengines. These features allow an increased capacity of congestedairports. The CFJ blended wing system hence will deliver superior energyefficiency and favorable impacts to environment as described.

Now referring to FIG. 25, in another configuration, fluid may bedispensed from a plurality of injection openings 34 along the span ofthe aircraft and recovered by one or more recovery openings 24 alsopositioned along the span of the aerodynamic structure 10. Moreover, theinjection and recovery openings may only span a portion of theaerodynamic structure, rather than the entire length. For example, theinjection opening 22 may be partially obstructed to achieve various jetsizes and various lift profiles. In particular, a plurality of tabs orflaps 36 may be included on the aerodynamic structure 10 proximate theinjection opening 22 to define the plurality of injection openings 34.The tabs 36 may be the same material as the aerodynamic structure 10 orairfoil, for example, steel or another alloy and may be any shape orsize. In an exemplary configuration, the tabs 36 have a substantiallyrounded or ovular cross-section to provide for improved airflow and aresized to obstruct a portion of the length of the injection opening 22.Each tab 36 may be substantially U-shaped such that the recessed portionof the “U” may engage and fit with the top portion of the aerodynamicstructure as shown in FIG. 26. The tabs 36 may further be welded orotherwise anchored to the aerodynamic structure. In a particularconfiguration as shown in FIG. 2, the tabs 36 are spaced along the spanof the airfoil. A jet is formed between two adjacent tabs 36 and airflowis blocked by the tabs 36 as each tab 36 is positioned to occlude aportion of the injection opening 24.

The arrangement and size of the tabs 36 may be variable depending on thedesired lift and thrust. For example, the table in Table 3 showsexemplary arrangements of tabs 36 along the length of the injectionopening 22. Different sizes and arrangements of tabs 36 are organized byrespective obstruction factors (OF). The OF is defined as the obstructedarea divided by the area of the unobstructed injection opening 22. Basedon Newtonian fluid flow, for a given mass flow rate, increasing the OFwill result in an increase in fluid flow velocity through the pluralityof openings 34 owing to decreased flow area. As a result, depending onthe OF, the number of openings 34, and the Angle of Attack (AoA),certain lift profiles may be achieved.

For example, in an exemplary configuration to test the various tabs 36configurations, the range of AoA was varied between 0° to 35°. Nominalfree stream velocity was V_(∞)=10 m/s for all tests and the chordReynolds number was about 195,000. The jet momentum coefficient Cμ wasdefined as

$C_{\mu} = \frac{\overset{.}{m}\; V_{jet}}{\frac{1}{2}\rho_{\infty}V_{\infty}^{2}S}$where {dot over (m)}is the jet mass flow rate, V_(jet) is the jetvelocity, ρ_(∞)is the free stream density, V_(∞)is the free streamvelocity, and S is the planform area of the airfoil. Cμ changes when OFis changed even if this kept constant. For comparison purposes, the jetmomentum coefficient for the open slot airfoil or CFJ (FIG. 1) describedabove was used and defined as:

$C_{\mu}^{*} = \frac{\overset{.}{m}\; V_{jet}^{*}}{\frac{1}{2}\rho_{\infty}V_{\infty}^{2}S}$where the superscript * stand for open slot airfoil described.

For a given OF, a large number of configurations can be obtaineddepending on the number of jet injection openings 34 and the openingsizes. Table 3 shows the list of the OF presented here, ranging from ⅕to¾, and the openingsizes. For each OF, two configurations were testedcontaining the larger (hereafter labeled A) and smaller number ofdiscrete tabs 36 (hereafter labeled B). Mass flow rates of {dot over(m)}=0 kg/s, 0.030 kg/s, 0.045 kg/s and 0.060 kg/s are used.Corresponding open slot airfoil momentum coefficients are Cμ*=0, 0.08,0.16 and 0.30 , respectively. Table 4 shows the values of Cμ and C_(μ)*and the corresponding jet exit velocity for all configurations having aplurality of injection openings 34. The following drag reductionefficiency parameter is defined as:

$\eta = {\frac{C_{D_{baseline}} - C_{D_{DCFJ}}}{C_{\mu}}.}$This quantity describes the efficiency of drag reduction at the energycost to pump the jet with the momentum coefficient of Cμ. In otherwords, it is a measure of how much injection jet momentum is absorbed asthe drag reduction.

TABLE 3 Matrix of discrete CFJ configuration # of Hole width Schematicrepresentation (openings are Name OF Config jets mm (% cord) injectionholes, solid lines are tabs) DCFJ 1/5 1/5 A 5 94.5 (16.0%)

B 2 236.2 (40.0%)

DCFJ 1/3 1/3 A 10 43.7 (7.4%)

B 5 98.4 (16.7%)

DCFJ 1/2 1/2 A 19 15.5 (2.6%)

B 3 98.4 (16.7%)

DCFJ 2/3 2/3 A 19 10.4 (1.76%))

B 9 21.9 (3.7%)

DCFJ 3/4 3/4 A 14 10.5 (1.79%))

B 5 29.5 (5.0%)

TABLE 4 Corresponding Cμ and C_(μ)* for discrete CFJ configurations Cμand V_(jet) (Discrete CFJ) {dot over (m)} (kg/s) C_(μ)* and V_(jet)(Open Slot) DCFJ 1/5 DCFJ 1/3 DCFJ 1/2 DCFJ 2/3 DCFJ 3/4 0.030 0.08 0.110.13 0.17 0.23 0.34 (25 m/s) (29 m/s) (38 m/s) (52 m/s) (73 m/s) (106m/s)) 0.045 0.16 0.21 0.23 0.30 0.49 0.67 (33 m/s) (43 m/s) (51 m/s) (69m/s) (109 m/s) (153 m/s) 0.060 0.30 0.36 0.41 0.58 0.89 1.32 (46 m/s)(56 m/s) (69 m/s) (97 m/s) (150 m/s) (231 m/s)

As explained below, “DCFJ” stands for discrete co-flow jet, whichincludes configurations with a plurality of tabs 36 creating theplurality of injection openings 34. The configurations with highernumber of injection openings 34 (A) systematically generate more liftthan the alternate configurations (B). This is because configuration Awith smaller injection openings 34 generates higher jet velocity, higherentrainment, stronger jet mixing, and thus higher circulation. Thistrend increases as the value of C_(μ)* is increased. Overall, a liftincrease between 10% and 40% over the open slot airfoil can be obtainedusing configuration A instead of B. For increased OF, the differencebetween the two configurations seems to decrease with increasing AoA.These results suggest that smaller sized injection openings 34 withhigher jet velocity are more effective to enhance lift.

Compared to open slot airfoils, DCFJ ⅕(meaning ⅕of the injection opening22 is obstructed) and DCFJ ⅓show a small decrease in lift for all C_(μ)*and all AoA except past the stall angle (AoA≅25°). For DCFJ ½, a clearincrease in lift can be observed for all AoA. This trend increases forincreasing values of C_(μ)*. Finally, DCFJ ⅔and DCFJ ¾show a substantiallift increase over the open slot CFJ for all AoA and all C_(μ)*.

Evaluation of the lift performance increase for DCFJs compared tobaseline and open slot CFJ are shown in Fig. Increasing C_(μ)* and OFsystematically provides an increase in lift. The difference between DCFJ⅔and DCFJ ¾is not obvious and that, for lift enhancement, they appearapproximately equivalent. Compared to the open slot CFJ, the DCFJs showsimprovement for OF higher than ½. While DCFJ ½show only a 10% increasein lift, DCFJ ⅔and DCFJ ¾show a 30% to 50% increase. These results areconsiderable considering the magnitude of lift achievable and the energyexpenditure of the CFJ pump. For example, using the DCFJ ⅔at C_(μ)*=0.08provides comparable lift coefficient with the open slot CFJ that needstwice the flow rate. For open slot CFJ, the maximum lift is increased by1.5 to 1.8 times. For DCFJ ¾, the maximum lift is increased by 2.73times using the same mass flow rate as the open slot CFJ airfoil.

Similarly to the lift results, DCFJ ⅕and DCFJ ⅓show only a smalldecrease in drag compared to the open slot CFJ. Negative values of dragare observed with increased range of AoA. For OF larger than ½, asignificant drag reduction is observed for all C_(μ)*. A largedifference between DCFJ ⅔and DCFJ ¾is observed: while being comparablefor lift increase, DCFJ ¾shows lower drag than DCFJ ⅔. Negative valuesof drag are observed for AoA up to 30°. DCFJs are seen to provideincreased values of thrust, with up to 4 times of the thrust of the openslot CFJ airfoil. These results are significant since this negativedrag/thrust component can effectively reduce the overall drag of anaircraft or helicopter rotor blade, providing a significant reduction inthe engine power required to move the aircraft or rotate the rotorblades.

All DCFJ airfoils (with OF>½) show net lift augmentation thatintensifies with increasing values of C_(μ)* and grows linearly withAoA. The drag reduction is nearly constant at different AoA for a fixedC_(μ)*. Similar to the lift augmentation, the higher the value of C_(μ)*is, the larger the drag is reduced. Open slot CFJ results range from 40%to 70%, depending on the value of C_(μ)*. For OF>½, the drag reductionefficiency is found to reach higher value, approaching 80% for all AoAand all C_(μ). This means that the energy cost to generate injection jetmomentum not only benefits the system with a significant lift gain, butalso transfers nearly 80% of jet momentum to drag reduction. Hence theoverall energy expenditure is small compared to the efficiency gainmeasured as the ratio of lift to drag L/D.

In addition to the enhancement of lift and drag reduction, DCFJ airfoilsalso significantly increase stall AoA. For DCFJ ⅔and DCFJ ¾, stall angleis approximately 30°, a 30% increase from the baseline airfoil stallangle of 23°. Higher stall AoA gives more stall margin and increases thesafety of aircraft.

In summary, the configurations with higher number injection openings 34and smaller tab 36 sizes show significantly better performanceenhancement for all flow rates. Compared to an open slot CFJ airfoilusing the same mass flow rate, a discrete CFJ airfoil can provide anadditional 50% increase in maximum lift, 30% stall AoA increase, and300% drag reduction. The DCFJ airfoil with more injection openings 34and large obstruction factor, namely DCFJ ⅔and DCF ¾, provide thegreatest performance enhancement. While the lift in both cases iscomparable, the DCFJ ¾supplies an additional 10% increase in thrustcompared to DCFJ ⅔. The pumping pressure ratio and power consumption areincreased by reducing the injection jet size and increasing the numberof openings 34. For all the tests shown here, the pumping pressure ratioremained under 1.2. The DCFJ power consumption is about 10 times higherthan that for open slot CFJ, and is up to 600 W. Nearly 80% of theinjection momentum is translated to drag reduction, which indicates thata CFJ airfoil is highly energy efficient.

It will be appreciated by persons skilled in the art that the presentinvention is not limited to what has been particularly shown anddescribed herein above. In addition, unless mention was made above tothe contrary, it should be noted that all of the accompanying drawingsare not to scale. A variety of modifications and variations are possiblein light of the above teachings without departing from the scope andspirit of the invention, which is limited only by the following claims.

What is claimed is:
 1. An aerodynamic structure, comprising: a leadingedge; a trailing edge; an injection opening proximate the leading edgehaving a length; a plurality of distinct tabs positioned within theinjection opening, the plurality of distinct tabs partially obstructingthe injection opening to define a plurality of injection openings; atleast one recovery opening located between the plurality of injectionopenings and the trailing edge; and one or more engine portionspositioned between the recovery opening and the trailing edge.
 2. Theaerodynamic structure according to claim 1, wherein the plurality oftabs are spaced evenly along the length of the injection opening.
 3. Theaerodynamic structure according to claim 1, the plurality of tabsincluding at least five tabs.
 4. The aerodynamic structure according toclaim 1, wherein the plurality of tabs are substantially the same size.5. The aerodynamic structure according to claim 1, wherein half of thelength of the injection opening is obstructed.
 6. The aerodynamicstructure according to claim 1, wherein two-thirds of the length ofinjection opening is obstructed.
 7. The aerodynamic structure accordingto claim 1, wherein the aerodynamic structure defines a flying wing. 8.The aerodynamic structure according to claim 1, further comprising anequal number of the plurality of injection openings and recoveryopenings.
 9. The aerodynamic structure according to claim 1, furthercomprising a wing having a span, and wherein the recovery openingextends along substantially the entire length of the span.
 10. Theaerodynamic structure according to claim 9, wherein the length of theinjection opening is substantially equal to the length of the span. 11.The aerodynamic structure according to claim 1, wherein the one or moreengine portions provide a pressurized fluid output to the injectionopening.
 12. An aircraft, comprising: an aircraft body defining aleading edge, a trailing edge, a first wing, and a second wing; aninjection opening having a length proximate the leading edge; aplurality of flaps disposed proximate the injection opening, theplurality of flaps at least partially obstructing the injection openingto define a plurality of injection openings; at least one recoveryopening located between the injection opening and the trailing edge; andone or more engines defining a fluid intake in fluid communication withthe recovery opening.
 13. The aircraft according to claim 12, whereinthe plurality of flaps obstruct half the length of the injectionopening.
 14. The aircraft according to claim 12, wherein the pluralityof flaps obstruct two-thirds of the length of the injection opening. 15.The aircraft according to claim 12, wherein the recovery opening is influid communication with a fluid path leading to the injection opening.16. The aircraft according to claim 12, wherein the one or more enginesprovide a pressurized fluid output to the injection opening.
 17. Theaircraft according to claim 12, wherein the plurality of flaps includesat least five flaps.
 18. The aircraft according to claim 12, furthercomprising an equal number of the plurality of injection openings andrecovery openings.
 19. An aerodynamic structure, comprising: a leadingedge, a trailing edge, a first wing, and a second wing, each of thefirst and second wings defining an injection opening; a plurality ofdistinct tabs mounted within the injection openings of the first andsecond wings, the plurality of distinct tabs defining a plurality ofinjection openings spanning the length of each of the first and secondwings; at least one recovery opening located between the plurality ofinjection openings and the trailing edge; and one or more engineportions positioned between the recovery opening and the trailing edge,the one or more engine portions in fluid communication with theplurality of injection openings and the recovery opening.